Turbofan gas turbine engines (which may be referred to simply as ‘turbofans’) are typically employed to power aircraft. Turbofans are particularly useful on commercial aircraft where fuel consumption is a primary concern. Typically a turbofan gas turbine engine will comprise an axial fan driven by an engine core. The engine core is generally made up of one or more turbines which drive respective compressors via coaxial shafts. The fan is usually driven directly off an additional lower pressure turbine in the engine core.
The fan comprises an array of radially extending fan blades mounted on a rotor and will usually provide, in current high bypass gas turbine engines, around seventy-five percent of the overall thrust generated by the gas turbine engine. The remaining portion of air from the fan is ingested by the engine core and is further compressed, combusted, accelerated and exhausted through a nozzle. The engine core exhaust mixes with the remaining portion of relatively high-volume, low-velocity air bypassing the engine core through a bypass duct.
To satisfy regulatory requirements, such engines are required to demonstrate that if part or all of a fan blade were to become detached from the remainder of the fan, that the detached parts are suitably captured within the engine containment system.
The fan is radially surrounded by a fan casing. It is known to provide the fan casing with a fan track liner and a containment system designed to contain any released blades or associated debris. Often, the fan track liner can form part of the fan containment system.
The fan track liner typically includes an annular layer of abradable material which surrounds the fan blades. During operation of the engine, the fan blades rotate freely within the fan track liner. At maximum speed the blades may cut a path into this abradable layer creating a seal against the fan casing and minimising air leakage around the blade tips during cruise. Further incursions can occur during gusts or take off rotations over time.
A conventional fan containment system or arrangement 100 is illustrated in FIG. 1 and surrounds a fan comprising an array of radially extending fan blades 40. Each fan blade 40 has a leading edge 44, a trailing edge 45 and fan blade tip 42. The fan containment arrangement 100 comprises a fan case 150. The fan case 150 has a generally frustoconical or cylindrical annular casing element 152 and a hook 154. The hook 154 is positioned axially forward of an array of radially extending fan blades 40. A fan track liner 156 is mechanically fixed or directly bonded to the fan case 150. The fan track liner 156 is provided as a structural intermediate to bridge a deliberate gap provided between the fan case 150 and the fan blade tip 42.
The fan track liner 156 has, in circumferential layers, an attrition liner 158 (also referred to as an abradable liner or an abradable layer), an intermediate layer which in this example is a honeycomb layer 160, and a septum 162. The septum layer 162 acts as a bonding, separation, and load spreading layer between the attrition liner 158 and the honeycomb layer 160. The honeycomb layer 160 may be an aluminium honeycomb. The tips 42 of the fan blades 40 are intended to pass as close as possible to the attrition liner 158 when rotating. The attrition liner 158 is therefore designed to be abraded away by the fan blade tips 42 during abnormal operational movements of the fan blade 40 and to just touch during the extreme of normal operation to ensure the gap between the rotating fan blade tips 42 and the fan track liner 156 is as small as possible without wearing a trench in the attrition liner 158. During normal operations of the gas turbine engine, ordinary and expected movements of the fan blade 40 rotational envelope cause abrasion of the attrition liner 158. This allows the best possible seal between the fan blades 40 and the fan track liner 156 and so improves the effectiveness of the fan in driving air through the engine.
The purpose of the hook 154 is to ensure that, in the event that a fan blade 40 detaches from the rotor of the fan 12, the fan blade 40 will not be ejected through the front, or intake, of the gas turbine engine. During such a fan-blade-off event, the fan blade 40 is held by the hook 154 and a trailing blade (not shown) forces the held released blade rearwards where the released blade is contained. Thus the fan blade 40 is unable to cause damage to structures outside of the gas turbine engine casings.
As can be seen from FIG. 1, for the hook 154 to function effectively, a released fan blade 40 must penetrate the attrition liner 158 in order for its forward trajectory to intercept with the hook. If the attrition liner 158 is too hard then the released fan blade 40 may not sufficiently crush the fan track liner 156.
However, the fan track liner 156 must also be stiff enough to withstand the rigours of normal operation without sustaining damage. This means the fan track liner 156 must be strong enough to withstand ice and other foreign object impacts without exhibiting damage for example. Thus there is a design conflict, where on one hand the fan track liner 156 must be hard enough to remain undamaged during normal operation, for example when subjected to ice impacts, and on the other hand allow the tip 42 of the fan blade 40 to penetrate the attrition liner 158. It is a problem of balance in making the fan track liner 156 sufficiently hard enough to sustain foreign object impact, whilst at the same time, not be so hard as to alter the preferred hook-interception trajectory of a fan blade 40 released from the rotor. Ice that impacts the fan casing rearwards of the blade position is resisted by a reinforced rearward portion 164 of the fan track liner.
An alternative fan containment system is indicated generally at 200 in FIG. 2. The fan containment system 200 includes a fan track liner 256 that is connected to the fan casing 250 at both an axially forward position and an axially rearward position. At the axially forward position, the fan track liner is connected to the casing at hook 254 via a fastener 266 that is configured to fail at a predetermined load. In the event of a fan blade detaching from the remainder of the fan, the fan blade impacts the fan track liner 256, the fastener 266 fails and the fan track liner pivots about a rearward point on the fan track liner. Such an arrangement is often referred to as a trap door arrangement. The trap door arrangement has been found to help balance the requirements for stiffness of the fan track liner with the requirements for resistance of operational impacts (e.g. ice impacts) ensuring a detached blade is held within the engine.
Often the fan track liner is formed from a plurality of arcuate fan track liner panels. Forming the fan track liner from a plurality of panels means that if a region of the fan track liner is damaged, only the affected panels need to be replaced. To ease removal of the fan track liner panels for such repair work, it is preferable for the fan track liner panels to be releasably connected to the fan casing, e.g. using bolts, instead of being adhered or bonded to the fan casing. However, when the fan track liner panels are releasably connected to the casing the fan track liner panels can vibrate during normal use, e.g. due to the pressure profile formed by passing blades during the operation of the fan. It is desirable to limit any such vibration to a minimal level. Excessive vibration can result in increased noise, increased blade to fan track liner clearance, and loss in engine performance. In very extreme cases, the vibration could lead to failure of the fan track liner panels or damaging interactions with the fan blades.
Several proposals for energy absorption during a fan blade off scenario are known, but these do not address the problem of damping vibration of the fan track liner during normal operation of a gas turbine engine.